High fan tip speed engine

ABSTRACT

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of rotatable fan blades, each fan blade defining a fan tip speed; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; and a gear box, wherein the turbomachine is operably coupled to the fan through the gear box, wherein a gear ratio of the gear box is greater than or equal to 1.2 and less than or equal to 3.0; wherein during operation of the turbofan engine at a rated speed the fan tip speed is greater than or equal to 1000 feet per second. In exemplary embodiments, during operation of the turbofan engine at the rated speed the fan pressure ratio is less than or equal to about 1.5.

FIELD

The present subject matter relates generally to a gas turbine engine, ormore particularly to a gas turbine engine configured to operate in amore efficient manner.

BACKGROUND

A turbofan engine generally includes a fan having a plurality of fanblades and a turbomachine arranged in flow communication with oneanother. Additionally, the turbomachine of the turbofan engine generallyincludes, in serial flow order, a compressor section, a combustionsection, a turbine section, and an exhaust section. In operation, air isprovided from the fan to an inlet of the compressor section where one ormore axial compressors progressively compress the air until it reachesthe combustion section. Fuel is mixed with the compressed air and burnedwithin the combustion section to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the turbine section and is then routed through the exhaustsection, e.g., to atmosphere.

However, efficiency losses in an upper span of the fan blades may resultin a less efficient turbofan engine.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to an exemplary embodiment of the present subjectmatter.

FIG. 2 is a close-up, schematic, cross-sectional view of a forward endof the exemplary gas turbine engine of FIG. 1 according to an exemplaryembodiment of the present subject matter.

FIG. 3 is a schematic view of an inlet to the exemplary gas turbineengine of FIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to an exemplary embodiment of the present subjectmatter.

FIG. 4 it is a schematic view of an inlet to a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

FIG. 5 is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 6 is a schematic cross sectional view of an embodiment of theturbine section shown in FIG. 5 according to another exemplaryembodiment of the present subject matter.

FIG. 7 is a schematic cross-sectional view of an exemplary gas turbineengine according to another exemplary embodiment of the present subjectmatter.

FIG. 8 is a flow diagram of a method for operating a turbofan engine inaccordance with an exemplary aspect of the present disclosure.

Corresponding reference characters indicate corresponding partsthroughout the several views. The exemplifications set out hereinillustrate exemplary embodiments of the disclosure, and suchexemplifications are not to be construed as limiting the scope of thedisclosure in any manner.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of thedisclosure, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the disclosure.

The following description is provided to enable those skilled in the artto make and use the described embodiments contemplated for carrying outthe disclosure. Various modifications, equivalents, variations, andalternatives, however, will remain readily apparent to those skilled inthe art. Any and all such modifications, variations, equivalents, andalternatives are intended to fall within the scope of the presentdisclosure.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations. Additionally, unlessspecifically identified otherwise, all embodiments described hereinshould be considered exemplary.

For purposes of the description hereinafter, the terms “upper”, “lower”,“right”, “left”, “vertical”, “horizontal”, “top”, “bottom”, “lateral”,“longitudinal”, and derivatives thereof shall relate to the disclosureas it is oriented in the drawing figures. However, it is to beunderstood that the disclosure may assume various alternativevariations, except where expressly specified to the contrary. It is alsoto be understood that the specific devices illustrated in the attacheddrawings, and described in the following specification, are simplyexemplary embodiments of the disclosure. Hence, specific dimensions andother physical characteristics related to the embodiments disclosedherein are not to be considered as limiting.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine, with forward referring to a position closer to an engineinlet and aft referring to a position closer to an engine nozzle orexhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Additionally, the terms “low,” “high,” or their respective comparativedegrees (e.g., lower, higher, where applicable) each refer to relativespeeds or pressures within an engine, unless otherwise specified. Forexample, a “low-pressure turbine” operates at a pressure generally lowerthan a “high-pressure turbine.” Alternatively, unless otherwisespecified, the aforementioned terms may be understood in theirsuperlative degree. For example, a “low-pressure turbine” may refer tothe lowest maximum pressure turbine within a turbine section, and a“high-pressure turbine” may refer to the highest maximum pressureturbine within the turbine section. An engine of the present disclosuremay also include an intermediate pressure turbine, e.g., an enginehaving three spools.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin. These approximating margins may apply to asingle value, either or both endpoints defining numerical ranges, and/orthe margin for ranges between endpoints.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

As used herein, the term “fan pressure ratio” refers to a ratio of anair pressure immediately downstream of the fan blades if a fan duringoperation of the fan to an air pressure immediately upstream of the fanblades of the fan during operation of the fan.

As used herein, the term “rated speed” with reference to a turbofanengine refers to a maximum rotational speed that the turbofan engine mayachieve while operating properly. For example, the turbofan engine maybe operating at the rated speed during maximum load operations, such asduring takeoff operations.

Also as used herein, the term “fan tip speed” as defined by theplurality of fan blades of the fan refers to a linear speed of an outertip of a fan blade along a radial direction during operation of the fan.

The present disclosure is generally related to a turbofan engine havinga fan defining a fan pressure ratio and a turbomachine operably coupledto the fan for driving the fan. The turbomachine defines a core airflowpath therethrough. Additionally, the fan of the turbofan engineincludes a plurality of rotatable fan blades each defining a fan tipspeed. The turbofan engine of the present disclosure also includes agear box, wherein the turbomachine is operably coupled to the fanthrough the gear box, wherein a gear ratio of the gear box is greaterthan or equal to 1.2 and less than or equal to 3.0. Furthermore, duringoperation of the turbofan engine at a rated speed the fan tip speed isgreater than or equal to 1000 feet per second. In an exemplaryembodiment, during operation of the turbofan engine at the rated speedthe fan pressure ratio is less than or equal to about 1.5.

It should be appreciated that the exemplary turbofan engine describedoperates contrary to conventional teachings. Specifically, conventionalengine operation teachings generally attempt to minimize fan speed withsuch gear ratios. For example, conventional engine operation teach lowfan speeds for better efficiencies. However, the turbofan enginedescribed herein operates contrary to these teachings by reducing thegear ratio while operating at a high fan tip speed, while maintaining arelatively low fan pressure ratio. The turbofan engine of the presentdisclosure achieves improved system efficiencies by enabling higher fantip speeds at lower fan pressure ratios. Furthermore, the turbofanengine of the present disclosure also provides pre-swirling flow forwardof the fan blade tip as described herein. Such may facilitate operationof the turbofan engine at the relatively high fan tip speeds withoutcreating undesirably high losses at the outer ends of the fan blades(e.g., as a result of the airflow over the fan blades separating andgenerating a turbulence in the airflow).

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1 , the gas turbine engine is an aeronautical,turbofan jet engine 10, referred to herein as “turbofan engine 10”,configured to be mounted to an aircraft, such as in an under-wingconfiguration or tail-mounted configuration. As shown in FIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference), a radial directionR, and a circumferential direction (i.e., a direction extending aboutthe axial direction A; not depicted). In general, the turbofan 10includes a fan section 14 and a turbomachine 16 disposed downstream fromthe fan section 14 (the turbomachine 16 sometimes also, oralternatively, referred to as a “core turbine engine”).

The exemplary turbomachine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a first, booster or low pressure (LP) compressor 22and a second, high pressure (HP) compressor 24; a combustion section 26;a turbine section including a first, high pressure (HP) turbine 28 and asecond, low pressure (LP) turbine 30; and a jet exhaust nozzle section32. A high pressure (HP) shaft 34 drivingly connects the HP turbine 28to the HP compressor 24. A low pressure (LP) shaft 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. The compressor section,combustion section 26, turbine section, and jet exhaust nozzle section32 are arranged in serial flow order and together define a core airflowpath 37 through the turbomachine 16. It is also contemplated thatthe present disclosure is compatible with an engine having anintermediate pressure turbine, e.g., an engine having three spools.

Referring still the embodiment of FIG. 1 , the fan section 14 includes avariable pitch, single stage fan 38, the turbomachine 16 operablycoupled to the fan 38 for driving the fan 38. The fan 38 includes aplurality of rotatable fan blades 40 coupled to a disk 42 in a spacedapart manner. As depicted, the fan blades 40 extend outwardly from disk42 generally along the radial direction R. Each fan blade 40 isrotatable relative to the disk 42 about a pitch axis P by virtue of thefan blades 40 being operatively coupled to a suitable actuation member44 configured to collectively vary the pitch of the fan blades 40, e.g.,in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal centerline 12 by LP shaft 36across a power gear box 46. The power gear box 46 includes a pluralityof gears for stepping down the rotational speed of the LP shaft 36 to amore efficient rotational fan speed. Accordingly, for the embodimentdepicted, the turbomachine 16 is operably coupled to the fan 38 throughthe power gear box 46.

In exemplary embodiments, the fan section 14 includes twenty-two (22) orfewer fan blades 40. In certain exemplary embodiments, the fan section14 includes twenty (20) or fewer fan blades 40. In certain exemplaryembodiments, the fan section 14 includes eighteen (18) or fewer fanblades 40. In certain exemplary embodiments, the fan section 14 includessixteen (16) or fewer fan blades 40. In certain exemplary embodiments,it is contemplated that the fan section 14 includes other number of fanblades 40 for a particular application.

During operation of the turbofan engine 10, the fan 38 defines a fanpressure ratio and the plurality of fan blades 40 each define a fan tipspeed. As will be described in greater detail below, the exemplaryturbofan engine 10 depicted defines a relatively high fan tip speed andrelatively low fan pressure ratio during operation of the turbofanengine at a rated speed. As used herein, the term “fan pressure ratio”refers to a ratio of an air pressure immediately downstream of the fanblades 40 during operation of the fan 38 to an air pressure immediatelyupstream of the fan blades 40 during operation of the fan 38. For theembodiment depicted in FIG. 1 , the fan 38 of the turbofan engine 10defines a relatively low fan pressure ratio. For example, the turbofanengine 10 depicted defines a fan pressure ratio less than or equal toabout 1.5. For example, in certain exemplary embodiments, the turbofanengine 10 may define a fan pressure ratio less than or equal to about1.4. The fan pressure ratio may be the fan pressure ratio of the fan 38during operation of the turbofan engine 10, such as during operation ofthe turbofan engine 10 at a rated speed.

As used herein, the term “rated speed” with reference to the turbofanengine 10 refers to a maximum rotational speed that the turbofan engine10 may achieve while operating properly. For example, the turbofanengine 10 may be operating at the rated speed during maximum loadoperations, such as during takeoff operations.

Also as used herein, the term “fan tip speed” defined by the pluralityof fan blades 40 refers to a linear speed of an outer tip of a fan blade40 along the radial direction R during operation of the fan 38. Inexemplary embodiments, the turbofan engine 10 of the present disclosurecauses the fan blades 40 of the fan 38 to rotate at a relatively highrotational speed. For example, during operation of the turbofan engine10 at the rated speed, the fan tip speed of each of the plurality of fanblades 40 is greater than or equal to 1000 feet per second and less thanor equal to 2250 feet per second. In certain exemplary embodiments,during operation of the turbofan engine 10 at the rated speed, the fantip speed of each of the fan blades 40 may be greater than or equal to1,250 feet per second and less than or equal to 2250 feet per second. Incertain exemplary embodiments, during operation of the turbofan engine10 at the rated speed, the fan tip speed of each of the fan blades 40may be greater than or equal to about 1,350 feet per second, such asgreater than about 1,450 feet per second, such as greater than about1,550 feet per second, and less than or equal to 2250 feet per second.

Referring still to the exemplary embodiment of FIG. 1 , the disk 42 iscovered by rotatable front nacelle or hub 48 aerodynamically contouredto promote an airflow through the plurality of fan blades 40.Additionally, the exemplary fan section 14 includes an annular fancasing or outer nacelle 50 that at least partially, and for theembodiment depicted, circumferentially, surrounds the fan 38 and atleast a portion of the turbomachine 16.

More specifically, the nacelle 50 includes an inner wall 52 and adownstream section 54 of the inner wall 52 of the nacelle 50 extendsover an outer portion of the turbomachine 16 so as to define a bypassairflow passage 56 therebetween. Additionally, for the embodimentdepicted, the nacelle 50 is supported relative to the turbomachine 16 bya plurality of circumferentially spaced outlet guide vanes 55.

During operation of the turbofan engine 10, a volume of air 58 entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58 as indicated by arrows 62 is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58 as indicated by arrow 64 is directed or routed into the core airflowpath 37. The ratio between an amount of airflow through the bypasspassage 56 (i.e., the first portion of air indicated by arrows 62) to anamount of airflow through the core air flowpath 37 (i.e., the secondportion of air indicated by arrows 64) is known as a bypass ratio.

In exemplary embodiments, the bypass ratio during operation of theturbofan engine 10 (e.g., at a rated speed) is less than or equal toabout eleven (11). For example, the bypass ratio during operation of theturbofan engine 10 (e.g., at a rated speed) may be less than or equal toabout ten (10), such as less than or equal to about nine (9).Additionally, the bypass ratio may be at least about two (2).

In other exemplary embodiments, the bypass ratio may generally bebetween about 7:1 and about 20:1, such as between about 10:1 and about18:1. The pressure of the second portion of air 64 is then increased asit is routed through the high pressure (HP) compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66.

In exemplary embodiments, a gear ratio of the gear box 46 is greaterthan or equal to 1.2 and less than or equal to 3.0. In some exemplaryembodiments, the gear ratio of the gear box 46 is greater than or equalto 1.2 and less than or equal to 2.6. In other exemplary embodiments,the gear ratio of the gear box 46 is greater than or equal to 1.2 andless than or equal to 2.0.

It should be appreciated that the exemplary turbofan engine describedoperates contrary to conventional teachings. Specifically, conventionalengine operation teachings generally attempt to minimize fan speed withsuch gear ratios. For example, conventional engine operation teach lowfan speeds for better efficiencies. However, the turbofan engine 10described herein operates contrary to these teachings by reducing thegear ratio while operating at a high fan tip speed, while maintaining arelatively low fan pressure ratio. The turbofan engine of the presentdisclosure achieves improved system efficiencies by enabling higher fantip speeds at lower fan pressure ratios. Furthermore, the turbofanengine of the present disclosure also provides pre-swirling flow forwardof the fan blade tip as described herein.

Referring still to FIG. 1 , the compressed second portion of air 64 fromthe compressor section mixes with fuel and is burned within thecombustion section to provide combustion gases 66. The combustion gases66 are routed from the combustion section 26, through the HP turbine 28where a portion of thermal and/or kinetic energy from the combustiongases 66 is extracted via sequential stages of HP turbine stator vanes68 that are coupled to the outer casing 18 and HP turbine rotor blades70 that are coupled to the HP shaft 34, thus causing the HP shaft 34 torotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft 36, thus causing the LP shaft 36 torotate, thereby supporting operation of the LP compressor 22 and/orrotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the turbomachine 16 to provide propulsive thrust.

Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the turbomachine 16.

In some exemplary embodiments, it will be appreciated that the exemplaryturbofan engine 10 of the present disclosure may be a relatively largepower class turbofan engine 10. Accordingly, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate a relativelylarge amount of thrust. More specifically, when operated at the ratedspeed, the turbofan engine 10 may be configured to generate at leastabout 20,000 pounds of thrust, such as at least about 25,000 pounds ofthrust, such as at least about 30,000 pounds of thrust. Accordingly, theturbofan engine 10 may be referred to as a relatively large power classgas turbine engine.

Moreover, it should be appreciated that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, the turbofan engine 10 may have any othersuitable configuration. For example, in certain exemplary embodiments,the fan may not be a variable pitch fan. Additionally, or alternatively,aspects of the present disclosure may be utilized with any othersuitable aeronautical gas turbine engine, such as a turboshaft engine,turboprop engine, turbojet engine, etc.

As discussed above, the turbofan engine 10 of the present disclosurealso provides pre-swirling flow forward of the fan blade tip. Referringnow also to FIG. 2 , a close-up, cross-sectional view of the fan section14 and forward end of the turbomachine 16 of the exemplary turbofanengine 10 of FIG. 1 is provided. In exemplary embodiments, the turbofanengine 10 includes an inlet pre-swirl feature located upstream of theplurality of fan blades 40 of the fan 38 and attached to or integratedinto the nacelle 50. More specifically, for the embodiment of FIGS. 1and 2 , the inlet pre-swirl feature is configured as a plurality of partspan inlet guide vanes 100. The plurality of part span inlet guide vanes100 are each cantilevered from of the outer nacelle 50 (such as from theinner wall 52 of the outer nacelle 50) at a location forward of theplurality of fan blades 40 of the fan 38 along the axial direction A andaft of the inlet 60 of the nacelle 50. More specifically, each of theplurality of part span inlet guide vanes 100 define an outer end 102along the radial direction R, and are attached to/connected to the outernacelle 50 at the radially outer end 102 through a suitable connectionmeans (not shown). For example, each of the plurality of part span inletguide vanes 100 may be bolted to the inner wall 52 of the outer nacelle50 at the outer end 102, welded to the inner wall 52 of the outernacelle 50 at the outer end 102, or attached to the outer nacelle 50 inany other suitable manner at the outer end 102.

Further, for the embodiment depicted, the plurality of part span inletguide vanes 100 extend generally along the radial direction R from theouter end 102 to an inner end 104 (i.e., an inner end 104 along theradial direction R). Moreover, as will be appreciated, for theembodiment depicted, each of the plurality of part span inlet guidevanes 100 are unconnected with an adjacent part span inlet guide vane100 at the respective inner ends 104 (i.e., adjacent part span inletguide vanes 100 do not contact one another at the radially inner ends104, and do not include any intermediate connection members at theradially inner ends 104, such as a connection ring, strut, etc.). Morespecifically, for the embodiment depicted, each part span inlet guidevane 100 is completely supported by a connection to the outer nacelle 50at the respective outer end 102 (and not through any structureextending, e.g., between adjacent part span inlet guide vanes 100 at alocation inward of the outer end 102 along the radial direction R). Aswill be discussed below, such may reduce an amount of turbulencegenerated by the part span inlet guide vanes 100.

Moreover, is depicted, each of the plurality of part span inlet guidevanes 100 do not extend completely between the outer nacelle 50 and,e.g., the hub 48 of the turbofan engine 10. More specifically, for theembodiment depicted, each of the plurality of inlet guide vane define aninlet guide vane (“IGV”) span 106 along the radial direction R, andfurther each of the plurality of part span inlet guide vanes 100 furtherdefine a leading edge 108 and a trailing edge 110. The IGV span 106refers to a measure along the radial direction R between the outer end102 and the inner end 104 of the part span inlet guide vane 100 at theleading edge 108 of the part span inlet guide vane 100. Similarly, itwill be appreciated, that the plurality of fan blades 40 of the fan 38define a fan blade span 112 along the radial direction R. Morespecifically, each of the plurality of fan blades 40 of the fan 38 alsodefines a leading edge 114 and a trailing edge 116, and the IGV span 106refers to a measure along the radial direction R between a radiallyouter tip and a base of the fan blade 40 at the leading edge 114 of therespective fan blade 40.

For the embodiment depicted, the IGV span 106 is at least about fivepercent of the fan blade span 112 and up to about fifty-five percent ofthe fan blade span 112. For example, in certain exemplary embodiments,the IGV span 106 may be between about fifteen percent of the fan bladespan 112 and about forty-five percent of the fan blade span 112, such asbetween about thirty percent of the fan blade span 112 and about fortypercent of the fan blade span 112.

Reference will now also be made to FIG. 3 , providing an axial view ofthe inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 . As will beappreciated, for the embodiment depicted, the plurality of part spaninlet guide vanes 100 of the turbofan engine 10 includes a relativelylarge number of part span inlet guide vanes 100. More specifically, forthe embodiment depicted, the plurality of part span inlet guide vanes100 includes between about twenty part span inlet guide vanes 100 andabout fifty part span inlet guide vanes 100. More specifically, for theembodiment depicted, the plurality of part span inlet guide vanes 100includes between about thirty part span inlet guide vanes 100 and aboutforty-five part span inlet guide vanes 100, and more specifically,still, the embodiment depicted includes thirty-two part span inlet guidevanes 100. Additionally, for the embodiment depicted, each of theplurality of part span inlet guide vanes 100 are spaced substantiallyevenly along the circumferential direction C. More specifically, each ofthe plurality of part span inlet guide vanes 100 defines acircumferential spacing 118 with an adjacent part span inlet guide vane100, with the circumferential spacing 118 being substantially equalbetween each adjacent part span inlet guide vane 100.

Although not depicted, in certain exemplary embodiments, the number ofpart span inlet guide vanes 100 may be substantially equal to the numberof fan blades 40 of the fan 38 of the turbofan engine 10. In otherembodiments, however, the number of part span inlet guide vanes 100 maybe greater than the number of fan blades 40 of the fan 38 of theturbofan engine 10, or alternatively, may be less than the number of fanblades 40 of the fan 38 of the turbofan engine 10.

Further, should be appreciated, that in other exemplary embodiments, theturbofan engine 10 may include any other suitable number of part spaninlet guide vanes 100 and/or circumferential spacing 118 of the partspan inlet guide vanes 100. For example, referring now briefly to FIG. 4, an axial view of an inlet 60 to a turbofan engine 10 in accordancewith another exemplary embodiment of the present disclosure is provided.For the embodiment of FIG. 4 , the turbofan engine 10 includes less thantwenty part span inlet guide vanes 100. More specifically, for theembodiment of FIG. 4 , the turbofan engine 10 includes at least eightpart span inlet guide vanes 100, or more specifically includes exactlyeight part span inlet guide vanes 100. Additionally, for the embodimentof FIG. 4 , the plurality of part span inlet guide vanes 100 are notsubstantially evenly spaced along the circumferential direction C. Forexample, at least certain of the plurality of part span inlet guidevanes 100 define a first circumferential spacing 118A, while other ofthe plurality of part span inlet guide vanes 100 define a secondcircumferential spacing 118B. For the embodiment depicted, the firstcircumferential spacing 118A is at least about twenty percent greaterthan the second circumferential spacing 118B, such as at least abouttwenty-five percent greater such as at least about thirty percentgreater, such as up to about two hundred percent greater. Notably, aswill be described in greater detail below, the circumferential spacing118 refers to a mean circumferential spacing between adjacent part spaninlet guide vanes 100. The non-uniform circumferential spacing may,e.g., offset structure upstream of the part span inlet guide vanes 100.

Referring back to FIG. 2 , it will be appreciated that each of theplurality of part span inlet guide vanes 100 is configured to pre-swirlan airflow 58 provided through the inlet 60 of the nacelle 50, upstreamof the plurality of fan blades 40 of the fan 38. As briefly discussedabove, pre-swirling the airflow 58 provided through the inlet 60 of thenacelle 50 prior to such airflow 58 reaching the plurality of fan blades40 of the fan 38 may reduce separation losses and/or shock losses,allowing the fan 38 to operate with the relatively high fan tip speedsdescribed above with less losses of in efficiency.

Referring to FIGS. 5 and 6 , in another exemplary embodiment, a turbofanengine 210 of the present disclosure includes a turbine section 290having a vaneless counter rotating turbine.

The embodiment illustrated in FIGS. 5 and 6 includes similar componentsto the embodiment illustrated in FIGS. 1-4 . For the sake of brevity,these similar components and the similar steps of using turbofan engine210 (FIGS. 5 and 6 ) will not all be discussed in conjunction with theembodiments illustrated in FIGS. 5 and 6 .

Referring to FIGS. 5 and 6 , the engine 210 has a longitudinal or axialcenterline 212 that extends there through for reference purposes. Theengine 210 defines an axial direction L, a radial direction R, anupstream end 299, and a downstream end 298 along the axial direction A.

Referring to FIGS. 5 and 6 the turbine section includes aninterdigitated turbine section 290. The engine 210 may include asubstantially tubular outer casing 218 that defines an annular inlet220. In the embodiment shown in FIG. 5 , the outer casing 218 encases orat least partially flows, in serial flow arrangement, a compressorsection 221, a combustion section 226, and the interdigitated turbinesection 290. The compressor section 221 defines a high pressurecompressor (HPC) 224 and an intermediate pressure compressor (IPC) 222in serial arrangement.

A fan assembly 214 is disposed forward or upstream 299 of the compressorsection 221. The fan assembly 214 includes a fan rotor 215. The fanrotor 215 includes one or more fan stages 241, in which each fan stage241 defines a plurality of blades 242 that are coupled to and extendoutwardly from the fan rotor 215 in the radial direction R. In anexemplary embodiment, the fan rotor 215 defines a single fan stage orsingle circumferentially adjacent arrangement of the plurality of blades242. In various other exemplary embodiments, the fan assembly 214 mayfurther define a plurality of the stages 241. The fan rotor 215 and fanblades 242 are together rotatable about the axial centerline 212. Anannular fan casing or nacelle 244 circumferentially surrounds at least aportion of the fan assembly 214 and/or at least a portion of the outercasing 218. In one embodiment, the nacelle 244 may be supported relativeto the outer casing 218 by a plurality of circumferentially-spacedoutlet guide vanes or struts 246. At least a portion of the nacelle 244may extend over an outer portion (in radial direction R) of the outercasing 218 so as to define a bypass airflow passage 248 therebetween.

As discussed above, the turbofan engine 210 of the present disclosurealso provides pre-swirling flow forward of the fan blade tip. Referringspecifically to FIG. 5 , the turbofan engine 210 also includes an inletpre-swirl feature located upstream of the plurality of fan blades 242and attached to or integrated into the nacelle 244. More specifically,for the embodiment of FIG. 5 , the inlet pre-swirl feature is configuredas a plurality of part span inlet guide vanes 100 as described abovewith respect to FIGS. 1-4 .

During operation of the engine 210, a volume of air as indicatedschematically by arrows 274 enters the engine 210 through an associatedinlet 276 of the nacelle and/or fan assembly 214. As the air indicatedby arrows 274 passes across the blades 242 of the fan assembly 214, aportion of the air as indicated schematically by arrows 278 is directedor routed into the bypass airflow passage 248 while another portion ofthe air as indicated schematically by arrows 280 is directed or throughthe fan assembly 214. The air indicated by arrows 280 is progressivelycompressed as it flows through the compressor section 221 toward thecombustion section 226.

The now compressed air, as indicated schematically by arrows 282, flowsinto the combustion section 226 where a fuel is introduced, mixed withat least a portion of the compressed air indicated by arrows 282, andignited to form combustion gases 286. The combustion gases 286 flow intothe turbine section 290, causing rotary members of the turbine section290 to rotate and support operation of respectively coupled rotarymembers in the compressor section 221 and/or fan assembly 214.

Referring now specifically to FIG. 6 , an exemplary embodiment of theturbine section 290 of the engine 210 is generally provided. The turbinesection 290 includes a first rotating component 310 interdigitated witha second rotating component 320 along the axial direction A. The firstrotating component 310 includes one or more connecting airfoils 316coupled to a radially extended rotor 315. The second rotating component320 includes an inner shroud 312 defining a plurality of inner shroudairfoils 317 extended outward of the inner shroud 312 along the radialdirection R. In various embodiments, the inner shroud 312 and/or theouter shroud 314 are formed or defined by a plurality of hubs, disks, ordrums defining an axial or longitudinal flowpath, such as a portion of acore flowpath 270 of compressed air 282 and combustion gases 286 throughthe engine 210 from the upstream end 299 to the downstream end 298.

In various embodiments, the first rotating component 310 furtherincludes an outer shroud 314 defining a plurality of outer shroudairfoils 318 extended inward of the outer shroud 314 along the radialdirection R. The outer shroud 318 may be coupled to the one or moreconnecting airfoils 316 and extended forward or upstream. The pluralityof outer shroud airfoils 318 may extend inward of the outer shroud 314in interdigitation with the plurality of inner shroud airfoils 317extended along the radial direction R from the inner shroud 312 of thesecond rotating component 320. In various embodiments, the secondrotating component 320 is disposed upstream of the one or moreconnecting airfoils 316 of the first rotating component 310 and ininterdigitation with the plurality of outer shroud airfoils 318 extendedfrom the first rotating component 310.

In one embodiment, the first and second rotating components 310, 320 maytogether define at least three stages of rotating airfoils (e.g.,connecting airfoil 316, second rotating component 320, and outer shroudairfoil 318 of first rotating component 310). In another embodiment, thefirst and second rotating components 310, 320 together define betweenthree and ten stages or rows of rotating airfoils.

The engine 210 further includes a gear assembly 245 within the turbinesection 290, such as inward along the radial direction R, or downstreamof the turbine section 290 along the axial direction A. For example, thegear assembly 245 may be disposed toward the downstream end 298 of theengine 210. As another example, the gear assembly 245 is disposeddownstream of the turbine section 290 within an exhaust frame 350 (e.g.,inward of the exhaust frame 250 along the radial direction R and alignedwith the exhaust frame 250 along the axial direction A). The gearassembly 245 includes an input accessory 247 and an output accessory249. A second shaft 321 is connected to the input accessory 247 andprovides power into the gear assembly 245. The second rotating component320 is coupled to the second shaft 321 and provides power into the gearassembly 245. The first rotating component 310 is coupled to the one ormore output accessories 249 of the gear assembly 245. The one or moreoutput accessories 249 rotate the first rotating component 310 about theaxial centerline 212 at a first speed. The second rotating component 320coupled to the second shaft 321 and rotates about the axial centerline212 at a second speed. In various embodiments, the second speed at whichthe second rotating component 320 rotates is greater than the firstspeed at which the first rotating component 310 rotates.

Referring still to FIG. 6 , the engine 210 further includes a firstshaft 236 extended in the axial direction A and through the gearassembly 245 from an upstream end of the gear assembly 245 to adownstream end of the gear assembly 245. The rotor 315 of the firstrotating component 310 is coupled to the first shaft 236. In exemplaryembodiments, the first shaft 236 may be directly connected to an innerrotor of the counter rotating turbine. In other exemplary embodiments,the first shaft 236 may be connected to an outer rotor of the counterrotating turbine. In various embodiments, the rotor 315 of the firstrotating component 310 is rotatably coupled to the one or more outputaccessories 249 of the gear assembly 245. In one embodiment, the rotor315 defines a housing 324 generally surrounding the gear assembly 245and coupled to the first shaft 236. In various embodiments, the housing324 includes an axial portion 326 and one or more radial portions 325,327. In one embodiment, a first radial portion 325 extends at leastpartially in the radial direction R from the rotor 315 to the upstreamend of the gear assembly 245. The first radial portion 325 is coupled tothe one or more output accessories 249 of the gear assembly 245. Inanother embodiment, a second radial portion 327 extends at leastpartially in the radial direction R from the downstream end of the gearassembly 245 to the first shaft 236. The second radial portion 327 iscoupled to the one or more output accessories 249 of the gear assembly245 toward the downstream end 298 of the gear assembly 245.

In one embodiment, such as shown in FIG. 6 , the axial portion 326 ofthe housing 324 may connect the first radial portion 325 and the secondradial portion 327 at least partially in the axial direction A. Invarious embodiments, the first radial portion 325, the second radialportion 327, and/or the axial portion 326 may each define asubstantially annular structure generally concentric about the axialcenterline 212.

In various embodiments, the second rotating component 320 may define aradially extended rotor portion 328 extended from the second shaft 321to the inner shroud 312. The rotor portion 328 of the second rotatingcomponent 320 is rotatably coupled to the second shaft 321. In variousembodiments, the inner shroud 312 and the rotor portion 328 may definean integral structure. In one embodiment, the inner shroud airfoil 317may further define an integral structure with the rotor portion 328 andinner shroud 312. In another embodiment, the rotor portion 328 defines ahub into which the plurality of inner shroud airfoils 317 installs.

The engine 210 shown and described in regard to FIG. 6 may define atorque path from the second rotating component 320 to the second shaft321, from the second shaft 321 to the input accessory 247 of the gearassembly 245, and from the one or more output accessories 249 of thegear assembly 245 to the housing 324, such as shown at the second radialportion 327 in FIG. 6 , of the first rotating component 310, and fromthe rotating component 310 to the first shaft 236. Still further, theengine 210 may define the torque path from the first rotating component310 to the first shaft 236 via the second radial portion 327 of thefirst rotating component 310. In one embodiment further including theaxial portion 326 of the first rotating component 310, the torque pathmay be defined from the first rotating component 310 through housing324, such as through the axial portion 326 to the second radial portion327, and to the first shaft 236.

In various embodiments, the first rotating component 310 rotates in afirst direction 361 and the second rotating component 320 rotates in asecond direction 362 opposite of the first direction 361. The firstrotating component 310, and the output accessory 249 of the gearassembly 245 to which the first rotating component 310 via the firstshaft 236, rotates in the first direction 361 as the second rotatingcomponent 320, coupled to the input accessory 247 of the gear assembly245 via the second shaft 321, rotates in the second direction 362. Inexemplary embodiments, the gear assembly 245 is configured as areversing reduction gear assembly.

In various embodiments, the gear assembly 245 defines a plurality ofgears in which the input accessory 247 rotates at a speed greater thanthe output accessory 249 or the first shaft 236 receiving power from thegear assembly 245. As such, the second rotating component 320 rotates ata speed greater than the first rotating component 310. Additionally, thesecond rotating component 320 rotates at a speed greater than the firstrotating component 310 in a direction opposite of the first rotatingcomponent 310.

In exemplary embodiments, a gear ratio of the gear assembly 245 isgreater than or equal to 1.2 and less than or equal to 3.0. In someexemplary embodiments, the gear ratio of the gear box 46 is greater thanor equal to 1.2 and less than or equal to 2.6. In other exemplaryembodiments, the gear ratio of the gear box 46 is greater than or equalto 1.2 and less than or equal to 2.0.

It should be appreciated that the exemplary turbofan engine describedoperates contrary to conventional teachings. Specifically, conventionalengine operation teachings generally attempt to minimize fan speed withsuch gear ratios. For example, conventional engine operation teach lowfan speeds for better efficiencies. However, the turbofan enginedescribed herein operates contrary to these teachings by reducing thegear ratio while operating at a high fan tip speed, while maintaining arelatively low fan pressure ratio. The turbofan engine of the presentdisclosure achieves improved system efficiencies by enabling higher fantip speeds at lower fan pressure ratios. Furthermore, the turbofanengine of the present disclosure also provides pre-swirling flow forwardof the fan blade tip as described herein.

In exemplary embodiments, the turbofan engine 210 of the presentdisclosure causes the fan blades 242 to rotate at a relatively highrotational speed. For example, during operation of the turbofan engine210 at the rated speed, the fan tip speed of each of the plurality offan blades 242 is greater than or equal to 1000 feet per second and lessthan or equal to 2250 feet per second. In certain exemplary embodiments,during operation of the turbofan engine 210 at the rated speed, the fantip speed of each of the fan blades 242 may be greater than or equal to1,250 feet per second and less than or equal to 2250 feet per second. Incertain exemplary embodiments, during operation of the turbofan engine210 at the rated speed, the fan tip speed of each of the fan blades 242may be greater than or equal to about 1,350 feet per second, such asgreater than about 1,450 feet per second, such as greater than about1,550 feet per second, and less than or equal to 2250 feet per second.

In exemplary embodiments, the fan 214 of the turbofan engine 210 definesa relatively low fan pressure ratio. For example, the turbofan engine210 depicted defines a fan pressure ratio less than or equal to about1.5. For example, in certain exemplary embodiments, the turbofan engine210 may define a fan pressure ratio less than or equal to about 1.4. Thefan pressure ratio may be the fan pressure ratio of the fan 214 duringoperation of the turbofan engine 210, such as during operation of theturbofan engine 210 at a rated speed.

In some exemplary embodiments, the gear assembly 245 may define acompound gearbox. In some exemplary embodiments, the gear assembly 245defines a reversing rotating compound gearbox.

Referring still to FIG. 6 , the engine 210 further includes the exhaustframe 350 disposed aft or downstream 298 of the first and secondrotating components 310, 320. The exhaust frame 350 includes one or moreexhaust vanes 352 extended in the radial direction R. The exhaust frame350 further includes a static support structure 354 extended inwardalong the radial direction R. The support structure 354 generallydefines a static annular casing defining one or more fasteninglocations. The gear assembly 245 is coupled to the exhaust frame 350 atthe support structure 354. In various embodiments, the gear assembly 245and the support structure 354 together transfer torque or power from thesecond rotating component 320 through the gear assembly 245 to the firstshaft 236 via the second radial portion 327 of the first rotatingcomponent 310.

In various embodiments, the exhaust frame 350 further includes a capcovering or concealing the gear assembly 245 within the exhaust frame350 from external view and environmental conditions. The cap may beremoved to provide relatively quick access to the gear assembly 245, thefirst shaft 236, or other components of the engine 210 with rear mountedgear assembly 245, in proximity to an unobstructed aft, outside portionof the engine 210, in contrast to a forward mounted gear assemblyconfiguration (e.g., within a fan assembly or low pressure compressor),in which the fan assembly is generally removed to access the gearassembly.

Referring still to FIG. 6 , in various embodiments, the first and secondrotating component 310, 320 together define a low pressure turbine (LPT)rotor. In such embodiments, the first shaft 236 defines a low pressure(LP) shaft connected and rotatable with the fan rotor 215 of the fanassembly 214. The fan assembly 214 is driven collectively by the firstrotating component 310 and the second rotating component 320. Byarranging the engine 210 such that the first rotating component 310 iscoupled directly to the first shaft 236 that is coupled to the fan rotor215, and by arranging the second rotating component 320 as coupled tothe gear assembly 245 that is coupled at the output accessory 249 to thefirst shaft 236, in one embodiment the first rotating component 310transmits approximately 25% to about 75% of power or torque for rotationof the fan assembly 214. In another embodiment, the second rotatingcomponent 320 transmits approximately 30% to about 60% of power ortorque for rotation of the fan assembly 214, in which the secondrotating component 320 transmits power or torque through the gearassembly 245 to the first shaft 236 to the fan assembly 214.Additionally, interdigitating the first and second rotating components310, 320 to define the LPT rotor results in efficiency and performancebenefits due to relatively large flowpath velocities, reduced airfoilcount (i.e., removed stationary vanes between rotating components),and/or reduced longitudinal dimensions of the LPT rotor.

Referring now back to FIG. 5 , the turbine section 290 further includesa third rotating component 330 disposed forward or upstream of the oneor more connecting airfoils 316 of the first rotating component 310. Thethird rotating component 330 includes a plurality of third airfoils 332extended outward along the radial direction R. In one embodiment, thethird rotating component 330 is disposed forward or upstream 299 of thefirst and second rotating component 310, 320.

In exemplary embodiments, the counter rotating turbine engine withreversing reduction gear assembly may further increase engine efficiencyand performance by providing a turbine and gear assembly arrangementthat permits a three-spool engine configuration.

It will be appreciated, however, that in other exemplary embodiments, atwo-spool engine configuration in which the low pressure compressor isdriven by the counter rotating turbine in addition to the fan may beprovided. Further, in other exemplary embodiments, the gear assembly 245may, e.g., be located forward of the combustion section 226. Moreover,although the embodiment of FIG. 6 depicts one or more bearings, in otherembodiments the engine may include any other configuration of bearingspositioned at any other suitable location.

Referring now to FIG. 7 , a schematic cross-sectional view of anexemplary gas turbine engine according to another exemplary embodimentof the present subject matter is provided. In the exemplary embodimentof FIG. 7 , a turbofan engine 410 is provided that includes adifferential gear box 456 that drives a fan 412 at a different speedthan a vaneless counter rotating turbine.

Referring to FIG. 7 , housed within engine frame 411 is an engine coreportion designated generally 424 which serves as the combustion gasgenerator means for engine 410. In an exemplary embodiment, core portion424 includes high pressure compressor 426, combustion chamber 428, andturbine assembly 430 for driving compressor 426, all depictedschematically in FIG. 7 .

For example, the differential gear box 456 is inserted between a turbineassembly 430, e.g., a counter rotating turbine, and the fan 412.Referring to FIG. 7 , the differential gear box 456 is driven by therotors of the turbine assembly 430, e.g., a counter rotating turbine,with the output connected to a shaft of the fan 412. This enables theturbine to operate at a very high speed, while keeping a low fanpressure ratio as described herein.

In exemplary embodiments, a gear ratio of the differential gear box 456is greater than or equal to 1.2 and less than or equal to 3.0. In someexemplary embodiments, the gear ratio of the differential gear box 456is greater than or equal to 1.2 and less than or equal to 2.6. In otherexemplary embodiments, the gear ratio of the differential gear box 456is greater than or equal to 1.2 and less than or equal to 2.0. Notably,for a gear box including two input power sources and a single outputpower source, the gear ratio may be measured as the ratio of arotational speed quickest input (in RPM) to a rotational speed of theoutput (also in RPM).

In exemplary embodiments, the turbine section 290 (FIGS. 5 and 6 ) orturbine assembly 430 (FIG. 7 ), e.g., a counter rotating turbine,includes one or more of the rotating or rotor components, e.g., a firstturbine rotor and a second turbine rotor, and the gear box comprises adifferential gear box 456 (FIG. 7 ), wherein the fan 214, 412 rotates ina same direction and at a same speed as the first turbine rotor, andwherein the second turbine rotor is geared via the differential gear box456. In such embodiments, one of the inputs is also directly tied to theoutput. For example, it is contemplated that a sun gear and a ring gearmay be tied together, and a planetary carrier may be the other input.

It should be appreciated that the exemplary turbofan engine describedoperates contrary to conventional teachings. Specifically, conventionalengine operation teachings generally attempt to minimize fan speed withsuch gear ratios. For example, conventional engine operation teach lowfan speeds for better efficiencies. However, the turbofan enginedescribed herein operates contrary to these teachings by reducing thegear ratio while operating at a high fan tip speed, while maintaining arelatively low fan pressure ratio. The turbofan engine of the presentdisclosure achieves improved system efficiencies by enabling higher fantip speeds at lower fan pressure ratios. Furthermore, the turbofanengine of the present disclosure also provides pre-swirling flow forwardof the fan blade tip as described herein.

In exemplary embodiments, the turbofan engine 410 of the presentdisclosure causes fan blades 414 of the fan 412 to rotate at arelatively high rotational speed. For example, during operation of theturbofan engine 410 at the rated speed, the fan tip speed of each of theplurality of fan blades 414 is greater than or equal to 1000 feet persecond and less than or equal to 2250 feet per second. In certainexemplary embodiments, during operation of the turbofan engine 410 atthe rated speed, the fan tip speed of each of the fan blades 414 may begreater than or equal to 1,250 feet per second and less than or equal to2250 feet per second. In certain exemplary embodiments, during operationof the turbofan engine 410 at the rated speed, the fan tip speed of eachof the fan blades 414 may be greater than or equal to about 1,350 feetper second, such as greater than about 1,450 feet per second, such asgreater than about 1,550 feet per second, and less than or equal to 2250feet per second.

In exemplary embodiments, the fan 412 of the turbofan engine 410 definesa relatively low fan pressure ratio. For example, the turbofan engine410 depicted defines a fan pressure ratio less than or equal to about1.5. For example, in certain exemplary embodiments, the turbofan engine410 may define a fan pressure ratio less than or equal to about 1.4. Thefan pressure ratio may be the fan pressure ratio of the fan 412 duringoperation of the turbofan engine 410, such as during operation of theturbofan engine 410 at a rated speed.

Referring now to FIG. 8 , a method 500 of operating a turbofan enginecomprising a fan, a turbomachine, and a gear box, wherein theturbomachine is operably coupled to the fan through the gear box isdepicted. The exemplary method 500 may be utilized to operate one ormore of the engines described above with reference to FIGS. 1 through 7.

For the exemplary aspect of FIG. 8 , the method 500 generally includesat (502) rotating a fan of a turbofan engine with a turbomachine suchthat the fan defines a fan pressure ratio that is less than or equal to1.5, and a fan blade of the fan defines a fan tip speed greater than orequal to 1,000 feet per second, and a gear ratio of a gear box isgreater than or equal to 1.2 and less than or equal to 3.0 as describedin detail above with reference to FIGS. 1 through 7 .

For the exemplary aspect depicted, rotating the fan of the turbofanengine with the turbomachine includes at (504) rotating the fan of theturbofan engine with the turbomachine such that the fan tip speed isgreater than or equal to 1250 feet per second.

For the exemplary aspect depicted, rotating the fan of the turbofanengine with the turbomachine includes at (506) rotating the fan of theturbofan engine with the turbomachine with the gear ratio of the gearbox greater than or equal to 1.2 and less than or equal to 2.6.

It should be appreciated that the exemplary turbofan engine describedoperates contrary to conventional teachings. Specifically, conventionalengine operation teachings generally attempt to minimize fan speed withsuch gear ratios.

For example, conventional engine operation teach low fan speeds forbetter efficiencies. However, the turbofan engine described hereinoperates contrary to these teachings by reducing the gear ratio whileoperating at a high fan tip speed, while maintaining a relatively lowfan pressure ratio. The turbofan engine of the present disclosureachieves improved system efficiencies by enabling higher fan tip speedsat lower fan pressure ratios. Furthermore, the turbofan engine of thepresent disclosure also provides pre-swirling flow forward of the fanblade tip as described herein.

Further aspects of the disclosure are provided by the subject matter ofthe following clauses:

1. A turbofan engine comprising: a fan comprising a plurality ofrotatable fan blades, each fan blade defining a fan tip speed; aturbomachine operably coupled to the fan for driving the fan, theturbomachine comprising a compressor section, a combustion section, anda turbine section in serial flow order and together defining a core airflowpath; and a gear box, wherein the turbomachine is operably coupledto the fan through the gear box, wherein a gear ratio of the gear box isgreater than or equal to 1.2 and less than or equal to 3.0; whereinduring operation of the turbofan engine at a rated speed the fan tipspeed is greater than or equal to 1000 feet per second.

2. The turbofan engine of any preceding clause, wherein the fan definesa fan pressure ratio during operation of the turbofan engine, andwherein during operation of the turbofan engine at the rated speed thefan pressure ratio is less than or equal to about 1.5.

3. The turbofan engine of any preceding clause, wherein during operationof the turbofan engine at the rated speed the fan tip speed is greaterthan or equal to 1250 feet per second and less than or equal to 2250feet per second.

4. The turbofan engine of any preceding clause, wherein the gear ratioof the gear box is greater than or equal to 1.2 and less than or equalto 2.6.

5. The turbofan engine of any preceding clause, wherein the gear ratioof the gear box is greater than or equal to 1.2 and less than or equalto 2.0.

6. The turbofan engine of any preceding clause, further comprising anouter nacelle at least partially surrounding the fan and theturbomachine, the outer nacelle defining a bypass passage with theturbomachine.

7. The turbofan engine of any preceding clause, further comprising apart span inlet guide vane extending from the outer nacelle at alocation forward of the plurality of fan blades of the fan along anaxial direction and aft of an inlet of the outer nacelle.

8. The turbofan engine of any preceding clause, wherein the part spaninlet guide vane is configured to pre-swirl an airflow provided throughthe inlet of the outer nacelle and upstream of the plurality of fanblades of the fan.

9. The turbofan engine of any preceding clause, wherein the turbinesection comprises a counter rotating turbine.

10. The turbofan engine of any preceding clause, wherein the gear boxcomprises a differential gear box that drives the fan at a differentspeed than the counter rotating turbine.

11. The turbofan engine of any preceding clause, wherein the counterrotating turbine includes a first turbine rotor and a second turbinerotor and the gear box comprises a differential gear box, wherein thefan rotates in a same direction and at a same speed as the first turbinerotor, and wherein the second turbine rotor is geared via thedifferential gear box.

12. A method of operating a turbofan engine comprising a fan, aturbomachine, and a gear box, wherein the turbomachine is operablycoupled to the fan through the gear box, the method comprising: rotatingthe fan of the turbofan engine with the turbomachine such that the fandefines a fan pressure ratio that is less than or equal to 1.5, and afan blade of the fan defines a fan tip speed greater than or equal to1,000 feet per second, and a gear ratio of the gear box is greater thanor equal to 1.2 and less than or equal to 3.0.

13. The method of any preceding clause, wherein rotating the fan of theturbofan engine with the turbomachine comprises rotating the fan of theturbofan engine with the turbomachine such that the fan tip speed isgreater than or equal to 1250 feet per second and less than or equal to2250 feet per second.

14. The method of any preceding clause, wherein rotating the fan of theturbofan engine with the turbomachine comprises rotating the fan of theturbofan engine with the turbomachine with the gear ratio of the gearbox greater than or equal to 1.2 and less than or equal to 2.6.

15. The method of any preceding clause, wherein rotating the fan of theturbofan engine with the turbomachine comprises rotating the fan of theturbofan engine with the turbomachine with the gear ratio of the gearbox greater than or equal to 1.2 and less than or equal to 2.0.

16. The method of any preceding clause, further comprising pre-swirlinga flow of air provided to the fan of the turbofan engine duringoperation of the turbofan engine.

17. The method of any preceding clause, wherein pre-swirling the flow ofair provided to the fan of the turbofan engine comprises pre-swirlingthe flow of air provided to the fan of the turbofan engine using aninlet pre-swirl feature located upstream of the fan blade of the fan andattached to or integrated into a nacelle of the turbofan engine.

18. The method of any preceding clause, wherein the turbomachinecomprises a compressor section, a combustion section, and a turbinesection in serial flow order and together defining a core air flowpath,and wherein the turbine section comprises a counter rotating turbine.

19. The method of any preceding clause, wherein the gear box comprises adifferential gear box that drives the fan at a different speed than thecounter rotating turbine.

20. The method of any preceding clause, wherein the counter rotatingturbine includes a first turbine rotor and a second turbine rotor andthe gear box comprises a differential gear box, wherein the fan rotatesin a same direction and at a same speed as the first turbine rotor, andwherein the second turbine rotor is geared via the differential gearbox.

This written description uses examples to disclose the disclosure,including the best mode, and also to enable any person skilled in theart to practice the disclosure, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they include structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal languages of the claims.

While this disclosure has been described as having exemplary designs,the present disclosure can be further modified within the scope of thisdisclosure. This application is therefore intended to cover anyvariations, uses, or adaptations of the disclosure using its generalprinciples. Further, this application is intended to cover suchdepartures from the present disclosure as come within known or customarypractice in the art to which this disclosure pertains and which fallwithin the limits of the appended claims.

1-20. (canceled)
 21. A turbofan engine comprising: a fan comprising aplurality of rotatable fan blades, each fan blade defining a fan tipspeed; a turbomachine operably coupled to the fan for driving the fan,the turbomachine comprising a compressor section, a combustion section,and a turbine section in serial flow order and together defining a coreair flowpath; a gear box, wherein the turbomachine is operably coupledto the fan through the gear box, wherein a gear ratio of the gear box isgreater than or equal to 1.2 and less than or equal to 3.0; an outernacelle at least partially surrounding the fan and the turbomachine, theouter nacelle defining a bypass passage with the turbomachine; and apart span inlet guide vane extending from the outer nacelle at alocation forward of the plurality of rotatable fan blades of the fanalong an axial direction and aft of an inlet of the outer nacelle,wherein during operation of the turbofan engine at a rated speed the fantip speed is greater than or equal to 1000 feet per second.
 22. Theturbofan engine of claim 21, wherein the fan defines a fan pressureratio during operation of the turbofan engine, and wherein duringoperation of the turbofan engine at the rated speed the fan pressureratio is less than or equal to about 1.5.
 23. The turbofan engine ofclaim 21, wherein the gear ratio of the gear box is greater than orequal to 1.2 and less than or equal to 2.6.
 24. The turbofan engine ofclaim 21, wherein the gear ratio of the gear box is greater than orequal to 1.2 and less than or equal to 2.0.
 25. The turbofan engine ofclaim 21, wherein the part span inlet guide vane is configured topre-swirl an airflow provided through the inlet of the outer nacelle andupstream of the plurality of rotatable fan blades of the fan.
 26. Theturbofan engine of claim 21, wherein the turbine section comprises acounter rotating turbine.
 27. The turbofan engine of claim 26, whereinthe gear box comprises a differential gear box that drives the fan at adifferent speed than the counter rotating turbine.
 28. The turbofanengine of claim 26, wherein the counter rotating turbine includes afirst turbine rotor and a second turbine rotor and the gear boxcomprises a differential gear box, wherein the fan rotates in a samedirection and at a same speed as the first turbine rotor, and whereinthe second turbine rotor is geared via the differential gear box. 29.The turbofan engine of claim 21, wherein during operation of theturbofan engine at the rated speed the fan tip speed is greater than orequal to 1000 feet per second and less than or equal to 2250 feet persecond.
 30. The turbofan engine of claim 21, wherein during operation ofthe turbofan engine at the rated speed the fan tip speed is greater thanor equal to 1250 feet per second and less than or equal to 2250 feet persecond.
 31. The turbofan engine of claim 21, wherein during operation ofthe turbofan engine at the rated speed the fan tip speed is greater thanor equal to 1350 feet per second and less than or equal to 2250 feet persecond.
 32. A turbofan engine comprising: a fan comprising a pluralityof rotatable fan blades, each fan blade defining a fan tip speed; aturbomachine operably coupled to the fan for driving the fan; a gearbox, wherein the turbomachine is operably coupled to the fan through thegear box, wherein a gear ratio of the gear box is greater than or equalto 1.2 and less than or equal to 3.0; an outer nacelle at leastpartially surrounding the fan and the turbomachine; and a part spaninlet guide vane extending from the outer nacelle at a location forwardof the plurality of rotatable fan blades of the fan along an axialdirection and aft of an inlet of the outer nacelle, wherein the partspan inlet guide vane is configured to pre-swirl an airflow providedthrough the inlet of the outer nacelle and upstream of the plurality ofrotatable fan blades of the fan, and wherein during operation of theturbofan engine at a rated speed the fan tip speed is greater than orequal to 1000 feet per second.
 33. The turbofan engine of claim 32,wherein the fan defines a fan pressure ratio during operation of theturbofan engine, and wherein during operation of the turbofan engine atthe rated speed the fan pressure ratio is less than or equal to about1.5.
 34. The turbofan engine of claim 32, wherein the gear ratio of thegear box is greater than or equal to 1.2 and less than or equal to 2.6.35. The turbofan engine of claim 32, wherein the gear ratio of the gearbox is greater than or equal to 1.2 and less than or equal to 2.0. 36.The turbofan engine of claim 32, wherein the turbomachine includes aturbine section, and wherein the turbine section comprises a counterrotating turbine.
 37. The turbofan engine of claim 36, wherein the gearbox comprises a differential gear box that drives the fan at a differentspeed than the counter rotating turbine.
 38. The turbofan engine ofclaim 36, wherein the counter rotating turbine includes a first turbinerotor and a second turbine rotor and the gear box comprises adifferential gear box, wherein the fan rotates in a same direction andat a same speed as the first turbine rotor, and wherein the secondturbine rotor is geared via the differential gear box.
 39. The turbofanengine of claim 32, wherein during operation of the turbofan engine atthe rated speed the fan tip speed is greater than or equal to 1000 feetper second and less than or equal to 2250 feet per second.
 40. Theturbofan engine of claim 32, wherein during operation of the turbofanengine at the rated speed the fan tip speed is greater than or equal to1250 feet per second and less than or equal to 2250 feet per second.